Gearbox Assemblies with Idler Gears

This application claims the benefit of Italian Patent Application No. 102024000009487, filed on Apr. 24, 2024. The prior application is incorporated herein by reference in its entirety.

The present subject matter relates generally to gear assemblies and, in particular, to gear assembly arrangements suitable for reversing the direction of the rotational output of the gear assembly.

Gas turbine engines generally cause fan blades to rotate in the same direction on an aircraft. In some cases, it may be desirable to provide engines rotating in different directions. However, it is difficult to reverse the rotational direction of the fan blades of a gas turbine engine without significantly altering the design of the engine or the gearbox. Accordingly, there is a need for improvements in turbomachines to allow fan blades to rotate in different directions on the same aircraft.

Reference now will be made in detail to embodiments of the disclosure, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the disclosure, not limitation of the disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present disclosure without departing from the scope or spirit of the disclosure. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims and their equivalents.

The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations.

As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operational attitude of the gas turbine engine or vehicle. For example, with regard to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems.

Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

Disclosed herein are various embodiments of modifications to the gear assemblies of a rotor engine or turbofan engine. The gear assemblies disclosed herein can reverse the direction of rotation of one or more rotor engines. The gear assemblies disclosed herein can provide significant advantages over conventional systems.

For example, the gear assemblies disclosed herein may allow for the rotor engines of an aircraft to be rotated in opposite directions, reducing or eliminating one or more of several drawbacks associated with rotation of the engines in the same direction. For example, when an aircraft has engines rotating in the same direction, one of the engines may cast debris from operation towards the fuselage of the aircraft, or towards the other engine. This necessitates armoring the fuselage against possible impacts that may cause damage to one or more of the engines of the aircraft. Rotating the rotor engines in opposite direction, depending on engine location, may reduce the risk of damage from cross engine debris by causing the debris stream from all engines to flow away from other engines and from the fuselage of the aircraft. In turn, this may reduce the need to armor or shield portions of the aircraft against debris, allowing aircraft weight to be reduced, and avoid the risk of engine damage from cross-engine debris.

Rotating the rotor engines of the aircraft in opposite directions may additionally counteract the yaw forces which can occur when the aircraft engines are rotating in the same direction. The rotation of each engine can create a left or right yaw force depending on the direction in which the engines are rotating. When the engines are rotating in the same direction, these yaw forces can interact additively, causing a persistent yaw effect to the left or the right of the aircraft. When the engines are rotating in opposite directions, the yaw forces can partially or completely cancel each other out. With yaw forces eliminated or reduced, the need to provide a counteracting force from some other source may be eliminated or reduced in turn, which can improve aircraft performance and efficiency.

Furthermore, in the case of rotor engines located near the ends or tips of the wings of the aircraft, running the engines in opposite directions may allow both engines at the wing tips to be run in the inboard up rotational direction. This allows for control over the strength and direction of the wingtip vortex, which may result in improved wing efficiency.

Additionally, rotating the rotor engines of the aircraft in opposite directions may minimize aero-acoustic interactions in the cabin space in the aircraft. This may reduce the noise and discomfort caused to passengers by the operation of the rotor engines and improve the passenger experience.

By using alternative gearbox configurations, the direction of rotation can be changed prior to the booster and/or core flowpath, which minimizes the number of unique parts required, and minimizes part count, the need for retooling, the number of spare parts and modules that must be kept for engine repair, product cost, and maintenance cost. Additionally, product design flexibility may be improved.

Referring now to the drawings,is an exemplary embodiment of an engineincluding a gear assemblyaccording to aspects of the present disclosure. The engineincludes a fan assemblydriven by a core engine. In various embodiments, the core engineis a Brayton cycle system configured to drive the fan assembly. The core engineis shrouded, at least in part, by an outer casing. The fan assemblyincludes a plurality of fan blades. A vane assemblyis extended from the outer casing. The vane assemblyincluding a plurality of vanesis positioned in operable arrangement with the fan bladesto provide thrust, control thrust vector, abate or re-direct undesired acoustic noise, and/or otherwise desirably alter a flow of air relative to the fan blades. In some embodiments, the fan assemblyincludes between three (3) and twenty (20) fan blades. In particular embodiments, the fan assemblyincludes between ten (10) and sixteen (16) fan blades. In certain embodiments, the fan assemblyincludes twelve (12) fan blades. In certain embodiments, the vane assemblyincludes an equal or fewer quantity of vanesto fan blades.

In certain embodiments, such as depicted in, the vane assemblyis positioned downstream or aft of the fan assembly. However, it should be appreciated that in some embodiments, the vane assemblymay be positioned upstream or forward of the fan assembly. In still various embodiments, the enginemay include a first vane assembly positioned forward of the fan assemblyand a second vane assembly positioned aft of the fan assembly. The fan assemblymay be configured to desirably adjust pitch at one or more fan blades, such as to control thrust vector, abate or re-direct noise, and/or alter thrust output. The vane assemblymay be configured to desirably adjust pitch at one or more vanes, such as to control thrust vector, abate or re-direct noise, and/or alter thrust output. Pitch control mechanisms at one or both of the fan assemblyor the vane assemblymay co-operate to produce one or more desired effects described above.

The core engineis generally encased in outer casingdefining a maximum diameter. In certain embodiments, the engineincludes a length from a longitudinally forward endto a longitudinally aft end. In various embodiments, the enginedefines a ratio of length (L) to maximum diameter (Dmax) that provides for reduced installed drag. In one embodiment, L/Dmax is at least 2. In another embodiment, L/Dmax is at least 2.5. In some embodiments, the L/Dmax is less than 5, less than 4, and less than 3. In various embodiments, it should be appreciated that the L/Dmax is for a single unducted rotor engine.

The reduced installed drag may further provide for improved efficiency, such as improved specific fuel consumption. Additionally, or alternatively, the reduced drag may provide for cruise altitude engine and aircraft operation at or above Mach 0.5. In certain embodiments, the L/Dmax, the fan assembly, and/or the vane assemblyseparately or together configure, at least in part, the engineto operate at a maximum cruise altitude operating speed between approximately Mach 0.55 and approximately Mach 0.85.

Referring again to, the core engineextends in a radial direction R relative to an engine axis centerline. The gear assemblyreceives power or torque from the core enginethrough a power input source (e.g., input shaft) and provides power or torque to drive the fan assembly, in a circumferential direction C about the engine axis centerline, through a power output source (e.g., output shaft).

In certain embodiments, such as depicted in, the engineis an un-ducted thrust producing system, such that the plurality of fan bladesis unshrouded by a nacelle or fan casing. As such, in various embodiments, the enginemay be configured as an unshrouded turbofan engine, an open rotor engine, or a propfan engine. In particular embodiments, the engineis a single unducted rotor engine including a single row of fan blades. The engineconfigured as an open rotor engine includes the fan assemblyhaving large-diameter fan blades, such as may be suitable for high bypass ratios, high cruise speeds (e.g., comparable to aircraft with turbofan engines, or generally higher cruise speed than aircraft with turboprop engines), high cruise altitude (e.g., comparable to aircraft with turbofan engines, or generally higher cruise speed than aircraft with turboprop engines), and/or relatively low rotational speeds. Cruise altitude is generally an altitude at which an aircraft levels after climb and prior to descending to an approach flight phase. In various embodiments, the engine is applied to a vehicle with a cruise altitude up to approximately 65,000 ft. In certain embodiments, cruise altitude is between approximately 28,000 ft. and approximately 45,000 ft.

Although depicted above as an unshrouded or open rotor engine in, it should be appreciated that the gear assemblies disclosed herein may be applied to shrouded or ducted engines, partially ducted engines, aft-fan engines, or other turbomachine configurations, including those for marine, industrial, or aero-propulsion systems. In addition, the gear assemblies disclosed herein may also be applicable to turbofan, turboprop, or turboshaft engines.

For example,is a cross-sectional schematic illustration of an exemplary embodiment of an enginethat includes a gear assemblyin combination with a ducted fan propulsion system. However, unlike the open rotor configuration of, a fan assemblyand its fan bladesare contained within an annular fan caseand a vane assembly, including a plurality of vanes, extend radially between a fan cowland the inner surface of the fan case. As discussed above, the gear assemblies disclosed herein can provide for increased gear ratios for a fixed gear envelope (e.g., with the same size ring gear), or alternatively, a smaller diameter ring gear may be used to achieve the same gear ratios.

As shown in, a core engineis generally encased in an outer casing, and has a length extending from a longitudinally forward endto a longitudinally aft end. The exemplary core engine (for a ducted or unducted engine) can include a compressor section, a heat addition system(e.g., combustor), and an expansion sectiontogether in serial flow arrangement. The core engineextends circumferentially relative to an engine centerline axis. The core engineincludes a high-speed spool that includes a high-speed compressor and a high-speed turbine operably rotatably coupled together by a high-speed shaft. The heat addition systemis positioned between the high-speed compressor and the high-speed turbine. Various embodiments of the heat addition systeminclude a combustion section. The combustion section may be configured as a deflagrative combustion section, a rotating detonation combustion section, a pulse detonation combustion section, or other appropriate heat addition system. The heat addition systemmay be configured as one or more of a rich-burn system or a lean-burn system, or combinations thereof. In still various embodiments, the heat addition systemincludes an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof.

The core enginecan also include a booster or low-speed compressor positioned in flow relationship with the high-speed compressor. The low-speed compressor is rotatably coupled with the low-speed turbine via a low-speed shaftto enable the low-speed turbine to drive the low-speed compressor. The low-speed shaftis also operably connected to gear assemblyto provide power to the fan assemblyvia a power input source (e.g., input shaft), such as described further herein.

It should be appreciated that the terms “low” and “high”, or their respective comparative degrees (e.g., -er, where applicable), when used with compressor, turbine, shaft, or spool components, each refer to relative speeds within an engine unless otherwise specified. For example, a “low turbine” or “low-speed turbine” defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, lower than a “high turbine” or “high-speed turbine” at the engine. Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low turbine” or “low-speed turbine” may refer to the lowest maximum rotational speed turbine within a turbine section, a “low compressor” or “low speed compressor” may refer to the lowest maximum rotational speed turbine within a compressor section, a “high turbine” or “high-speed turbine” may refer to the highest maximum rotational speed turbine within the turbine section, and a “high compressor” or “high-speed compressor” may refer to the highest maximum rotational speed compressor within the compressor section. Similarly, the low-speed spool refers to a lower maximum rotational speed than the high-speed spool. It should further be appreciated that the terms “low” or “high” in such aforementioned regards may additionally, or alternatively, be understood as relative to minimum allowable speeds, or minimum or maximum allowable speeds relative to normal, desired, steady state, etc. operation of the engine.

As discussed in more detail below, the core engineincludes a gear assembly that is configured to transfer power from the expansion sectionand reduce an output rotational speed at the fan assemblyrelative to a low-speed turbine. Embodiments of the gear assemblies depicted and described herein can allow for gear ratios suitable for large-diameter unducted fans (e.g.,) or certain turbofans (e.g.,). Additionally, embodiments of the gear assemblies provided herein may be suitable within the radial or diametrical constraints of the core engine within the outer casing.

The gear assemblies described herein includes a gear set for decreasing the rotational speed of the fan assembly relative to the low speed (pressure) turbine. In operation, the rotating fan blades are driven by the low speed (pressure) turbine via gear assembly such that the fan blades rotate around the engine axis centerline and generate thrust to propel the engine, and hence an aircraft on which it is mounted, in the forward direction.

illustrates one or more layshaft pinsand compound planet gears,.illustrate the layshaft pins, compound planet gears,, and a sun gearwith a ring gear() and with a portion of the ring gear removed (). In the embodiments shown in, three compound planet gears are provided (,) and the ring gearcomprises two halves with an interconnecting flanged portion.also discloses a plurality of radial passages for oil scavenging.

illustrate a planet gear carrierwith a single compound planet gear (,) provided therein for clarity. The layshaft pinextends through an opening in the fore and aft sides of the planet gear carrier. In some embodiments, the carrier can be connected to the engine frame via a flexible support system, with the flexible support system being configured to collect oil and scavenge oil via holes at a lower portion.

In some embodiments, the gear ratio split between the first and second stages can range from 40% to 60% for each stage (i.e., from 40% to 60% for the first stage and from 60% to 40% for the second stage).

As discussed above, in some embodiments, the sun gear, planet gears,, and ring gearcan be double helical gears with first and second sets of helical teeth that are inclined at an acute angle relative to each other.

In the embodiment shown in, a gear assemblyis a star gear configuration in which the planet carrier is generally fixed (e.g., static) within the engine by support structure. The sun gearis driven by an input shaft(e.g., a low-speed shaft). A planet gear carrieris rotatably coupled to a layshaft of the compound planet gears,, and the ring gearis configured to rotate about a longitudinal engine axis centerlinein a circumferential direction, which in turn drives the power output source (e.g., a fan shaft) that is coupled to and configured to rotate with the ring gear to drive the fan assembly. In this embodiment, the low-speed shaftrotates in a circumferential direction that is the opposite of the direction in which a fan drive shaftrotates.

In other embodiments, the gear assembly can have a planetary configuration in which the ring gear is fixed (e.g., static) within the engine by a support structure. The sun gear is driven by an input shaft (i.e., low-speed shaft) and instead of the ring gear rotating, the planet carrier rotates in the same direction of the low-speed shaft rotation direction, to drive the power output source (e.g., a fan shaft) and the fan assembly.

Referring again to, the ring gearis coupled to the fan drive shaftto drive the fans. The sun gearis coupled to an input power source (e.g., input shaft). In some embodiments, the input shaft can be integrally formed with the sun gear. The bi-helical meshes of the planet gears axially balance the load over the four (phased) gear sets of each compound planet gear. The second stage of planet gearscan be supported by two rows of cylindrical roller bearingsat the planet bore. In addition, the fan drive shaftcan be supported by tapered roller bearings or angular ball bearings, which supports the fan drive shaftin an axially compact manner. In some embodiments, the roller bearingscan be formed from a ceramic material. In some embodiments, the roller bearingscan be lubricated by under-race lubrication, in which lubrication is directed under the inner race and forced out through a plurality of holes in the inner race. In some embodiments, as shown in, an inner supporting element of both sets of the roller bearingscan be a solid unique element.

In some embodiments, one of the pair of gear sets (e.g., one of the first and second gear sets, one of the third or fourth gear sets) is angularly clocked by a set amount of gear pitch relative to the other gear set. For example, the teeth of the first gear set can be angularly clocked by a first amount of the gear pitch relative to the teeth of the second gear set. The first amount can be between one fourth and one half. Similarly, the teeth of the third gear set can be angularly clocked by a second amount of the gear pitch relative to the teeth of the fourth gear set. The second amount can be between one fourth and one half.

The following are exemplary gear assemblies that can reverse the rotational direction of a turbofan engine according to the examples disclosed herein. In this way an aircraft can comprise at least one turbofan engine rotating in a first direction and at least one turbofan engine rotating in a second direction. For example, an aircraft with a pair of turbofan engines can include a first turbofan engine having fan blades rotating in a first rotational direction (e.g., clockwise or counterclockwise), and a second turbofan engine having fan blades rotating in a second rotating direction (e.g., clockwise or counterclockwise) that is opposite to the first rotational direction. For aircraft with more than two turbofan engines, the turbofan engines on the same side of the aircraft body can rotate in the same direction relative to each other, or different directions. Such assemblies may replace or be used with any of the gear assemblies previously described, and can be incorporated into any engine design, including those discussed above.

In one embodiment, the direction of a second turbofan engine of a pair of turbofan engines can be reversed while keeping a substantially similar configuration for the engine components by introducing a plurality of idler gears to the gear assembly driving the fan blades of the turbofan engine.show two exemplary gear assembly configurations for reversing the rotational direction of one turbofan engine of a pair of turbofan engines relative to the rotational direction of the other turbofan engine.

shows an epicyclic gear assemblyfor use in the first turbofan engine of the pair of turbofan engines similar to that illustrated in. Gear assemblycan be a star configuration with a ring gear, three compound planet gearshaving a first stageand a second stage, and a sun gear. In operation, sun gearis driven by an input shaft driven by the core engine output of the first engine in a first rotational direction (e.g., clockwise or counterclockwise). The sun gearengages with the first stageof the planet gear, causing the first stageand the second stageof the planet gear to rotate in a second rotational direction (e.g., counterclockwise or clockwise) that is opposite to the first rotational direction. The second stage of the planet gearengages with the ring gear, causing the ring gearto rotate in the second rotational direction as well (e.g., counterclockwise or clockwise). Ring gearis configured to drive the fan assembly of a turbofan engine such as turbofan engineorin the second rotational direction (e.g., clockwise or counterclockwise). In this way, the fan assembly of the first engine is driven in the opposite rotational direction of the input shaft from the core engine output of the first engine. While the example shown inshows a gear assembly with three compound planet gears, it should be understood that a smaller number of compound planet gears, such as two compound planet gears or one compound planet gear, or a larger number of planet gears, such as four, five, or six planet gears could also be used.

The gear assemblycan have a gear ratio between the input and output shafts that is from 5:1 to 14:1, from 6:1 to 12:1, from 7:1 to 11:1, or from 8:1 to 10:1. In certain specific examples, the gear assemblymay have a gear ratio of 5:1, 6:1, 7:1, 8:1, 9:1, 10:1, 11:1, 12:1, 13:1, 14:1, or any gear ratio in between. In one example, the gear assembly has a gear ratio of 8.7:1.

shows another epicyclic gear assemblysuitable for use with the second engine of the pair of turbofan engines. Gear assemblyhas a ring gear, three compound planet gearshaving a first stageand a second stage, a plurality of idler gears, and a sun gear. In operation, sun gearis driven by a turbine in a first rotational direction (e.g., clockwise or counterclockwise). The sun gearengages with the first stageof the planet gears, causing the first stageand the second stageof the planet gears to rotate in a second rotational direction (e.g., counterclockwise or clockwise) opposite to the first rotational direction. The second stageof the planet gears engages the idler gears, causing the idler gearsto rotate in the first rotational direction (e.g., clockwise or counterclockwise). The idler gearsengage the ring gear, causing the ring gearto rotate in the first rotational direction (e.g., clockwise or counterclockwise). Ring gearis configured to drive the rotating fan blades of a turbofan engine such as turbofan engineorin the first rotational direction (e.g., clockwise or counterclockwise). In this way, the fan assembly of the second engine is driven in the same rotational direction as the input shaft from the core engine output of the second engine. While the example shown inshows a gear assembly with three compound planet gears, it should be understood that a smaller number of compound planet gears, such as two compound planet gears or one compound planet gear, or a larger number of planet gears, such as four, five, or six planet gears could also be used. It should also be understood that a smaller number of idler gears, such as two idler gears or one idler gear, or a larger number of idler gears, such as four, five, or six idler gears could also be used.

Because the idler gearsare positioned between the second stageof the planet gearsand the ring gear, they may experience cyclical fatigue in two directions, compared with the one direction experienced by the planet gears. To address this additional direction of cyclical stress, in some examples, the ring gear, the planet gears, the idler gears, and the sun gear, and may be made with a greater gear module (i.e., with thicker teeth) to improve the expected service life of the part before failure necessitates repair or replacement.

Like gear assembly, gear assemblycan have a gear ratio between the input and output shafts that is from 5:1 to 14:1, from 6:1 to 12:1, from 7:1 to 11:1, or from 8:1 to 10:1. In certain specific examples, the gear assemblymay have a gear ratio of 5:1, 6:1, 7:1, 8:1, 9:1, 10:1, 11:1, 12:1, 13:1, 14:1, or any gear ratio in between. In one example, the gear assembly has a gear ratio of between 8.7:1 to 8.9:1. Preferably, the gear ratios of gear assemblyand gear assemblyare the same, or similar (e.g., within 5% of one another).

In this way, different turbofan engines (e.g. turbofan engines,) on the same aircraft can cause the respective fan assemblies to rotate in different directions. In addition, since the gear assemblies are similar except for the idler gears and related aspects, common components can be used in each of the two gear assemblies reducing the number of parts required to assemble and maintain the engines, and the two gear assemblies can achieve the same, or similar, outputs.

show another embodiment of two exemplary gear assembly configurations for reversing the rotational direction of one turbofan engine of a pair of turbofan engines relative to the rotational direction of the other turbofan engine. As in other embodiments, the arrangement ofadvantageously maintain a substantially similar configuration for the engine components between the two different arrangements of.

shows an epicyclic gear assemblyfor use in the first turbofan engine of the pair of turbofan engines similar to that illustrated in. Gear assemblycan be a star configuration with a ring gear, three compound planet gearshaving a first stageand a second stage, and a sun gear. In operation, sun gearis driven by an input shaft driven by the core engine output of the first engine in a first rotational direction (e.g., counterclockwise in this example). The sun gearengages with the first stageof the planet gear, causing the first stageand the second stageof the planet gear to rotate in a second rotational direction (e.g., clockwise in this example) that is opposite to the first rotational direction. The second stage of the planet gearengages with the ring gear, causing the ring gearto rotate in the second rotational direction as well (e.g., clockwise in this example). Ring gearis configured to drive the fan assembly of a turbofan engine such as turbofan engineorin the second rotational direction. In this way, the fan assembly of the first engine is driven in the opposite rotational direction of the input shaft from the core engine output of the first engine. While the example shown inshows a gear assembly with three compound planet gears, it should be understood that a smaller number of compound planet gears, such as two compound planet gears or one compound planet gear, or a larger number of planet gears, such as four, five, or six planet gears could also be used.

The gear assemblycan have a gear ratio between the input and output shafts that is from 5:1 to 14:1, from 6:1 to 12:1, from 7:1 to 11:1, or from 8:1 to 10:1. In certain specific examples, the gear assemblymay have a gear ratio of 5:1, 6:1, 7:1, 8:1, 9:1, 10:1, 11:1, 12:1, 13:1, 14:1, or any gear ratio in between. In one example, the gear assembly has a gear ratio of 8.7:1.